This invention relates to an arrangement for a gas turbine engine. More particularly, but not exclusively, this invention relates to a gas turbine engine rotor assembly and its surrounding casing.
A gas turbine engine comprises in series, a fan (or low pressure compressor), one or more higher pressure compressors, a combustion chamber for burning the compressed air from the compressors with fuel, and one or more turbines driven by the exhaust gases from the combustor. These components of the gas turbine engine are contained within an annular casing.
The compressors of gas turbine engines are commonly axial flow compressors which comprise alternate rows of rotating (rotor) blades and stationary (stator) vanes, to accelerate and diffuse the air passing therethrough until a required pressure rise is obtained. The casing surrounding the compressors normally comprises a number of axially adjacent cylindrical casing elements bolted together. The amount of clearance between the tips of the compressor rotor blades and the inner surface of the annular casing is an important consideration in gas turbine engine design as a large gap between the tips of the rotor blades and the inner casing surface would allow a larger amount of working fluid leakage over the tips of the blades than is desirable.
It is known to provide the inner surface of the compressor casing with an abradable lining which sacrificially wears away when rubbed by the tips of rotor blades. Compressor rotor blades are often manufactured from titanium, which can ignite when rubbed against an unlined inner surface of the casing. The provision of such an abradable lining assists in keeping the rotor tip clearance to a minimum and also helps to prevent rotor tip damage caused by titanium rotor tip fires.
It is also known to provide slots within the inner surface of the casing radially adjacent the compressor rotor blade tips. These slots provide the advantage of allowing recirculation of the working fluid over the blade tips, within the slots, which is known to provide aerodynamic benefits.
However, the provision of slots within the inner surface of the compressor casing, in combination with an abradable lining, is problematic. Machining such slots within a casing which has already been lined may damage the lining and reduce its benefit to engine efficiency. Applying the abradable lining after machining the slots, for example by blanking the slots prior to spraying the lining material, may result in ragged edges of lining material around the slots. Furthermore, lining material has a tendency to fall away from the casing.
If the inner surface of the compressor casing is not provided with an abradable lining, it is known to thin the tips of the compressor rotor blades. This encourages heat dissipation in the event that the blade tip rubs against the casing, thus reducing the risk of ignition of a titanium fire within the compressor. Blade damage under heavy rubbing may also be reduced. Alternatively, the blade tips may be hardened by the application of a hard coating which helps to reduce the risk of titanium fires by preventing the titanium rubbing directly against the material of the casing. As a further alternative, the blade tips may be cut back, so increasing the tip clearance. This reduces the likelihood of rubbing, but at a cost to efficiency.
According to the present invention there is provided an arrangement for a gas turbine engine, the arrangement comprising a rotor and a surrounding casing, said casing having a generally cylindrical inner surface, said inner surface having a first cylindrical section and a second, axially adjacent cylindrical section, said first section being provided with a plurality of recesses in a circumferential arrangement and said second section being provided with an abradable lining, said rotor comprising an annular array of rotor blades, each blade having an outer tip and each of said tips comprising a treated portion and an untreated portion wherein said treated portion of said tip is at least partly located opposite said first cylindrical section of the inner surface of said casing and said untreated portion of said tip is located opposite said second cylindrical section of the inner surface of said casing.
Preferably the first cylindrical section is disposed axially upstream from said second cylindrical section.
The recesses may comprise slots provided within the cylindrical inner surface of said casing. Each slot may be of generally rectangular cross-section.
The treated portion of the rotor blade tip may be provided at the leading edge region of the rotor blade.
The treated portion of said rotor blade tip may be harder than the untreated portion of said rotor blade tip.
The treated portion of said rotor blade tip may be provided with a hard coating, the coating being harder than the untreated portion of the blade tip. The hard coating may be an aluminium oxide.
The thickness of the treated portion of said rotor blade tip may be less than that of the untreated portion.
The rotor may be a high pressure compressor rotor for a gas turbine engine.
Also according to the present invention there is provided a gas turbine engine including a compressor incorporating an arrangement as described in any of the preceding nine paragraphs.